In a turbomachine, such as a gas turbine engine, air is pressurized in a compressor then mixed with fuel and burned in a combustor to generate hot combustion gases. The hot combustion gases are expanded within the turbine section where energy is extracted to power the compressor and to produce useful work, such as powering a propeller for an aircraft in flight or turning a generator to produce electricity. The hot combustion gas travels through a series of turbine stages. A turbine stage may include a row of stationary vanes followed by a row of rotating turbine blades, where the turbine blades extract energy from the hot combustion gas for powering the compressor and providing output power. Since the turbine blades are directly exposed to the hot combustion gas, they are typically provided with internal cooling circuits which channel a coolant, such as compressor bleed air, through the airfoil of the blade and through various film cooling holes around the surface thereof.
As turbine inlet temperatures increase and the pressure ratio across the turbine stage becomes higher, the cooling schemes for the first stage blades become more complicated. In particular, in order to enhance the cooling efficiency at the trailing edges of the blade airfoils, the cooling circuit configurations required for the trailing edges becomes increasingly intricate. As a consequence, the ceramic cores that are used to form the trailing edge cooling circuits have become more complex and fragile, with an associated complexity in the manufacturing process.
Accordingly, there is a need for a turbine blade airfoil design that is conducive to formation of intricate airfoil cooling circuits and which may provide an increase in the production yield during blade manufacture.